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Correlation equation relating creep deformation of tubes to stress and temperature functions, and application to design of heat exchanger tube
A procedure is described for sizing, scaling, positioning and performance loss calculation of a boundary layer trip protuberance. The theoretical results are compared with some experimental data. The application of boundary layer trips, such as secondary injection, wall angle discontinuities, circumferential grooves, and circumferential trip protuberances for reducing side loads on rocket nozzle walls is discussed.
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This report describes the Active Noise Cancellation (ANC) System designed by General Electric and tested in the NASA Lewis Research Center's (LERC) 48 inch Active Noise Control Fan (ANCF). The goal of this study is to assess the feasibility of using wall mounted secondary acoustic sources and sensors within the duct of a high bypass turbofan aircraft engine for global active noise cancellation of fan tones. The GE ANC system is based on a modal control approach. A known acoustic mode propagating in the fan duct is canceled using an array of flush-mounted compact sound sources. The canceling modal signal is generated by a modal controller. Inputs to the controller are signals from a shaft encoder and from a microphone array which senses the residual acoustic mode in the duct. The key results are that the (6,0) was completely eliminated at the 920 Hz design frequency and substantially reduced elsewhere. The total tone power was reduced 6.8 dB (out of a possible 9.8 dB). Farfield reductions of 15 dB (SPL) were obtained. The (4,0) and (4,1) modes were reduced simultaneously yielding a 15 dB PWL decrease. The results indicate that global attenuation of PWL at the target frequency was obtained in the aft quadrant using an ANC actuator and sensor system totally contained within the duct. The quality of the results depended on precise mode generation. High spillover into spurious modes generated by the ANC actuator array caused less than optimum levels of PWL reduction. The variation in spillover is believed to be due to calibration procedure, but must be confirmed in subsequent tests.
This report describes the Active Noise Cancellation System designed by General Electric and tested in the NASA Lewis Research Center's 48 inch Active Noise Control Fan. The goal of this study was to assess the feasibility of using wall mounted secondary acoustic sources and sensors within the duct of a high bypass turbofan aircraft engine for active noise cancellation of fan tones. The control system is based on a modal control approach. A known acoustic mode propagating in the fan duct is cancelled using an array of flush-mounted compact sound sources. Controller inputs are signals from a shaft encoder and a microphone array which senses the residual acoustic mode in the duct. The canceling modal signal is generated by a modal controller. The key results are that the (6,0) mode was completely eliminated at 920 Hz and substantially reduced elsewhere. The total tone power was reduced 9.4 dB. Farfield 2BPF SPL reductions of 13 dB were obtained. The (4,0) and (4,1) modes were reduced simultaneously yielding a 15 dB modal PWL decrease. Global attenuation of PWL was obtained using an actuator and sensor system totally contained within the duct.
The development of, and initial test data from, a nondetonating Pulse Detonation Engine (PDE) simulator tested in the NASA Glenn 1 x 1 foot Supersonic Wind Tunnel (SWT) is presented in this paper. The concept is a pulsed ejector driven by the simulated exhaust of a PDE. This pro- gram is applicable to a PDE entombed in a ramjet flowpath, i.e., a PDE combined-cycle propulsion system. The ejector primary flow is a pulsed, uiiderexpanded, supersonic nozzle simulating the supersonic waves ema- nating from a PDE, while the ejector secondary flow is the 1 x 1 foot SWT test section operated at subsonic Mach numbers. The objective is not to study the detonation details, but the wave physics including t,he start- ing vortices, the extent of propagation of the wave front, the reflection of the wave from the secondary flowpath walls, and the timing of these events of a pulsed ejector, and correlate these with Computational Fluid Dynamics (CFD) code predictions. Pulsed ejectors have been shown to result in a 3 to 1 improvement in LID (length-to-diameter) and a near 2 to 1 improvement in thrust augmentation over a steady ejector. This program will also explore the extent of upstream interactions between an inlet and large, periodically applied, backpressures to the inlet as would be present due to combustion tube detonations in a PDE. These interactions could result in inlet unstart or buzz for a supersonic mixed compression inlet. The design of the present experiment entailed the use of an 2-t diagram characteristics code to study the nozzle filling and purging timescales as well as a series of CFD analyses conducted using the WIND code. The WIND code is a general purpose CFD code for solution of the Reynolds averaged Navier-Stokes equations and can be applied to both steady state and time-accurate calculations. The first, proof-of-concept, test entry (spring 2001) pressure distributions shown here indicate the simulation concept was successful and therefore the experimental approach is sound.
Characteristics of continuous deposits near lunar craters larger than about 1 km wide are considered, and it is concluded that (1) concentric dunes, radial ridges, and braided lineations result from deposition of the collision products of ejecta from adjacent pairs of similarly oriented secondary-crater chains and are, therefore, concentrations of secondary-crater ejecta; (2) intracrater ridges are produced within preexisting craters surrounding a fresh primary crater by ricocheting and focusing of secondary-crater ejecta from the preexisting craters' walls; and (3) secondary cratering has produced many of the structures of the continuous deposits of relatively small lunar craters and is the dominant process for emplacement of most of the radial facies of the continuous deposits of large lunar craters and basins. The percentages of Imbrium ejecta in deposits and the nature of Imbrium sculpturing are investigated.
Thin-walled neoprene containers prevent secondary radiation, scatter, and undercut during radiographic inspection. The containers are filled with a mixture of barium sulfate, red lead, and petroleum jelly that achieves the required absorption rate.
The report describes a structure for a strong, lightweight corrugated sheet. The sheet is planar or curved and includes a plurality of corrugation segments, each segment being comprised of a generally U-shaped corrugation with a part-cylindrical crown and cap strip, and straight side walls and with secondary corrugations oriented at right angles to said side walls. The cap strip is bonded to the crown and the longitudinal edge of said cap strip extends beyond edge at the intersection between said crown and said side walls. The high strength relative to weight of the structure makes it desirable for use in aircraft or spacecraft.
The investigation was carried out using the rotating hot wire technique. Measurements were taken inside the end wall boundary layer to discern the effect of annulus and hub wall boundary layer, secondary flow, and tip leakage on the wake structure. Static pressure gradients across the wake were measured using a static stagnation pressure probe insensitive to flow direction changes. The axial and the tangential velocity defects, the radial component of velocity, and turbulence intensities were found to be very large as compared to the near and far wake regions. The radial velocities in the trailing edge region exhibited characteristics prevalent in a trailing vortex system. Flow near the blade tips found to be highly complex due to interaction of the end wall boundary layers, secondary flows, and tip leakage flow with the wake. The streamwise curvature was found to be appreciable near the blade trailing edge. Flow properties in the trailing edge region are quite different compared to that in the near and far wake regions with respect to their decay characteristics, similarity, etc. Fourier decomposition of the rotor wake revealed that for a normalized wake only the first three coefficients are dominant.
The periodic streamwise modulation of the supersonic and hypersonic boundary layers by a two dimensional first mode or second mode wave makes the resulting base flow susceptible to a broadband spanwise-periodic three dimensional type of instability. The principal parametric resonance of this instability (subharmonic) was analyzed using Floquet theory. The effect of Mach number and the effectiveness of wall cooling or wall suction in controlling the onset, the growth rate, and the vortical nature of the subharmonic secondary instability are assessed for both a first mode and a second mode primary wave. Results indicate that the secondary subharmonic instability of the insulated wall boundary layer is weakened as Mach number increases. Cooling of the wall destabilizes the secondary subharmonic of a second mode primary wave, but stabilizes it when the primary wave is a first mode. Suction stabilizes the secondary subharmonic at all Mach numbers.
The periodic streamwise modulation of the supersonic and hypersonic boundary layers by a two-dimensional first-mode or second-mode wave makes the resulting base flow susceptible to a broadband spanwise-periodic three-dimensional type of instability. The principal parametric resonance of this instability (subharmonic) has been analyzed using Floquet theory. The effect of Mach number and the effectiveness of wall cooling or wall suction in controlling the onset, the growth rate, and the vortical structure of the subharmonic secondary instability are assessed for both a first-mode and a second-mode primary wave. Results indicate that the secondary subharmonic instability of an insulated wall boundary layer is weakened as Mach number increases. Cooling of the wall destabilizes the secondary subharmonic of a second-mode primary wave, but stabilizes it when the primary wave is a first mode. Suction stabilizes the secondary subharmonic at all Mach numbers.
This paper reports the experimental study of the three-dimensional characteristics of the mean velocity of the rotor wake inside the annulus- and hub-wall boundary layers. The measurements were taken with a rotating three-sensor hot wire behind the rotor. This set of measurements probably represents the first set of comprehensive measurements taken inside the annulus- and hub-wall boundary layers. The wake was surveyed at several radial locations inside the boundary layer region and at several axial locations. Interaction of the wake with the annulus-wall boundary layer, secondary flow, tip-leakage flow, and the trailing vortex system results in slower decay and larger width of the wake. The presence of a strong vortex and its merger with the wake is also observed. The end-wall boundary layers and the secondary flow were found to have a substantial effect on both the decay characteristics and the profile of the wake. These and other measurements are reported and interpreted in this paper.
The original aim of this work was to shed some light on the physics of turbulence over rough walls using large-eddy simulations and the suboptimal-control wall boundary conditions introduced by Nicoud et al. It was hoped that, if that algorithm was used to fit the mean velocity profile of the simulations to that of a rough-walled channel, instead of to a smooth one, the wall stresses introduced by the control algorithm would give some indication of what aspects of rough walls are most responsible for the modification of the flow in real turbulence. It was similarly expected that the structure of the resulting velocity fluctuations would share some of the characteristics of rough-walled flows, thus again suggesting what is intrinsic and what is accidental in the effect of geometric wall roughness. A secondary goal was to study the effect of 'unphysical' boundary conditions on the outside flow by observing how a relatively major change of the target velocity profile, and therefore presumably of the applied wall stresses, modifies properties such as the dominant length scales of the velocity fluctuations away from the wall. As will be seen below, this secondary goal grew more important during the course of the study, which was carried out during a short summer visit of the first two authors to the CTR. It became clear that there are open questions about the way in which the control algorithm models the boundary conditions, even for smooth walls, and that these questions make the physical interpretation of the results difficult. Considerable more work in that area seems to be needed before even relatively advanced large-eddy simulations, such as these, can be used to draw conclusions about the physics of wall-bounded turbulent flows. The numerical method is the same as in Nicoud et al. The modifications introduced in the original code are briefly described in section 2, but the original paper should be consulted for a full description of the algorithm. The results are presented in section 3 and summarized in section 4. The elementary properties of turbulence over rough walls which are used in the text have been taken from recent reviews such as Raupach et al. or Jimenez.
A new dual-mode ramjet combustor concept intended for operation over a wide flight Mach number range is described. Subsonic combustion mode is similar to that of a traditional ram combustor which allows operation at higher efficiency, and to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle. The maximum flight Mach number of this scheme is governed largely by the same physics as its classical counterpart. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated. Given the parallel nature of the present scheme, overall flowpath length is less than that of present dual-mode configurations. Cycle analysis was done to define the flowpath geometry for computational fluid dynamics (CFD) analysis, and then to determine performance based on the CFD results. CFD results for Mach 5, 8, and 12 flight conditions indicate stable supersonic free-jet formation and nozzle reattachment, thereby establishing the basic feasibility of the concept. These results also reveal the structure of, and interactions between the free-jet and recirculating combustion chamber flows. Performance based on these CFD results is slightly less than that of the constant-pressure-combustion cycle analysis primarily due to these interactions. These differences are quantified and discussed. Additional CFD results at the Mach 8 flight condition show the effects of nozzle throat area variation on combustion chamber pressure, flow structure, and performance. Calculations with constant temperature walls were also done to evaluate heat flux and overall heat loads. Aspects of the concept that warrant further study are outlined. These include diffuser design, ramjet operation, mode transition, loss mechanisms, and the effects of secondary flow for wall cooling and combustion chamber pressurization. Also recommended is an examination of system-level aspects such as weight, thermal management and rocket integration as well as alternate geometries and variable geometry schemes.
Tissue-specific dirigent protein gene expression and associated dirigent (site) localization were examined in various organs of Forsythia intermedia using tissue printing, in situ mRNA hybridization and immunolabeling techniques, respectively. Dirigent protein gene expression was primarily noted in the undifferentiated cambial regions of stem sections, whereas dirigent protein sites were detected mainly in the vascular cambium and ray parenchyma cell initials. Immunolocalization also revealed cross-reactivity with particular regions of the lignified cell walls, these being coincident with the known sites of initiation of lignin deposition. These latter regions are considered to harbor contiguous arrays of dirigent (monomer binding) sites for initiation of lignin biopolymer assembly. Dirigent protein mRNA expression was also localized in the vascular regions of roots and petioles, whereas in leaves the dirigent sites were primarily associated with the palisade layers and the vascular bundle. That is, dirigent protein mediated lignan biosynthesis was initiated primarily in the cambium and ray cell initial regions of stems as well as in the leaf palisade layers, this being in accordance with the occurrence of the lignans for defense purposes. Within lignified secondary xylem cell walls, however, dirigent sites were primarily localized in the S(1) sublayer and compound middle lamella, these being coincident with previously established sites for initiation of macromolecular lignin biosynthesis. Once initiation occurs, lignification is proposed to continue through template polymerization.
This Minnowbrook IV 2003 workshop on Transition and Unsteady Aspects of Turbomachinery Flows includes the following topics: 1) Current Issues in Unsteady Turbomachinery Flows; 2) Global Instability and Control of Low-Pressure Turbine Flows; 3) Influence of End Wall Leakage on Secondary Flow Development in Axial Turbines; 4) Active and Passive Flow Control on Low Pressure Turbine Airfoils; 5) Experimental and Numerical Investigation of Transitional Flows as Affected by Passing Wakes; 6) Effects of Freestream Turbulence on Turbine Blade Heat Transfer; 7) Bypass Transition Via Continuous Modes and Unsteady Effects on Film Cooling; 8) High Frequency Surface Heat Flux Imaging of Bypass Transition; 9) Skin Friction and Heat Flux Oscillations in Upstream Moving Wave Packets; 10) Transition Mechanisms and Use of Surface Roughness to Enhance the Benefits of Wake Passing in LP Turbines; 11) Transient Growth Approach to Roughness-Induced Transition; 12) Roughness- and Freestream-Turbulence-Induced Transient Growth as a Bypass Transition Mechanism; 13) Receptivity Calculations as a Means to Predicting Transition; 14) On Streamwise Vortices in a Curved Wall Jet and Their Effect on the Mean Flow; 15) Plasma Actuators for Separation Control of Low Pressure Turbine Blades; 16) Boundary-Layer Separation Control Under Low-Pressure-Turbine Conditions Using Glow-Discharge Plasma Actuators; 17) Control of Separation for Low Pressure Turbine Blades: Numerical Simulation; 18) Effects of Elevated Free-Stream Turbulence on Active Control of a Separation Bubble; 19) Wakes, Calming and Transition Under Strong Adverse Pressure Gradients; 20) Transitional Bubble in Periodic Flow Phase Shift; 21) Modelling Spots: The Calmed Region, Pressure Gradient Effects and Background; 22) Modeling of Unsteady Transitional Flow on Axial Compressor Blades; 23) Challenges in Predicting Component Efficiencies in Turbomachines With Low Reynolds Number Blading; 24) Observations on the Causal Relationship Between Blade Count and Developing Rotating Stall in a Four Stage Axial Compressor; 25) Experimental and Numerical Study of Non-Linear Interactions in Transonic Nozzle Flow; 26) Clocking Effects on a Modern Stage and One-Half Transonic Turbine; 27) DNS and LES of Transition on Turbine Blades; 28) The Use of Cellular Automata in Modeling the Transition; 29) Predicting Unsteady Buffet Onset Using RANS Solutions; 30) Transition Modelling With the SST Turbulence Model and an Intermittency Transport; and 31) Equation Workshop Summary Transcript
The flow over wave shaped surfaces is studied using a Navier Stokes solver. Detailed comparisons with theoretical results are presented, including the stability of a laminar flow over wavy surfaces. Drag characteristics of nonplanar surfaces are predicted using the Navier-Stokes solver. The secondary instabilities of wall bounded and free shear flows are also discussed.